Electrically controlled solid rocket ignition system

ABSTRACT

An electrically controlled solid rocket ignition system including a resistance wire initiator for igniting pyrotechnic material. After ignition, a high voltage is applied between forward and aft electrodes which are positioned on or near the surface of the propellant. This produces an electrostatic field which permits controlled movement of heated ions and/or electrons from the flame zone to the remaining unignited surface of the propellant. Ignition delay time can thereby be precisely controlled.

Unite States Patent 1191 Turner 1 Jan. 23, 1973 i 1 ELECTRICALLYCONTROLLED SOLID ROCKET IGNITION SYSTEM [75] Inventor: Stewart W.Turner, Lancaster, Calif.

[73] Assignee: The United States of America as represented by theSecretary of the Air Force 22 Filed: March22, 1971 211 Appl.No.: 126,740

[52] U.S. Cl. ..l02/49.7, 102/70.2 A, 60/39.82 E,

[51] Int. Cl. ..F42c 19/12 [58] Field of Search ..l02/39, 49.7, 70.2 A,102;

[56] References Cited UNITED STATES PATENTS Brazil et al 260/256 x Rush..60/256 Bearer ..l02/l02 X Primary Examiner-Verlin R. PendegrassAttorney-Harry A. Herbert, Jr. and Ruth G. Codier [57] ABSTRACT Anelectrically controlled solid rocket ignition system including aresistance wire initiator for igniting pyrotechnic material. Afterignition, a high voltage is applied between forward and aft electrodeswhich are positioned on or near the surface of the propellant. Thisproduces an electrostatic field which permits controlled movement ofheated ions and/or electrons from the flame zone to the remainingunignited surface of the propellant. Ignition delay time can thereby beprecisely controlled.

1 Claim, 2 Drawing Figures ELECTRICALLY CONTROLLED SOLID ROCKET IGNITIONSYSTEM BACKGROUND OF THE INVENTION The present invention relates to asolid rocket ignition system, and more particularly to a system whichpermits command control of ignition delay time in a solid rocket motorby controlling the rate of flame spread over the surface of theunignited propellant.

In the standard pyrogen type igniter, hot gases are SUMMARY OF THEINVENTION The object of the present invention is, therefore, theprovision of an ignition device for solid rocket motors which will allowthe propulsion system a constant igni tion delay time, regardless of thetemperature of the rocket fuel.

A further object of the invention is to decrease the necessary weight ofthe missile by eliminating the need to mount a large power supply onboard the missile, and permit the power supply to be ground-mounted orto be positioned on a flying launch platform.

A further object of the invention is the provision of an ignition systemwherein the control of flame spread rates is so effective thatdifflculties usually experienced in slots and other such areas areeliminated, and flame spread proceeds.

The system of the invention consists of a resistance wire initiatorinstalled at the forward end of an internally ported solid rocket of anygiven design or internal geometry. The device of the invention iscapable of being modified. More sophisticated initiators can besubstituted for the resistance wire which is disclosed herein forpurposes of disclosure and exampliflcation. Connections for theresistance wire are made through an adaptor of high dielectric material,such as micarta (phenolic or epoxy impregnated linen). The resistancewire circuit is consumable and is vaporized immediately after thepyrotechnic material has ignited. A high voltage is then applied betweenforward and aft electrodes. The electrodes are positioned on or near thesurface of the propellant. An electrostatic field is thus producedbetween fore and aft electrodes, causing a force to be exerted oncertain charged particles (positive ions, negative ions and electrons)within the flame zone. The resulting motion of these charged particlesis primarily dependent upon magnitude and direction of the applied fieldforce, and can be controlled by external electrical circuitry. Movementof heated ions and electrons from the flame zone along the surface ofthe unignited propellant which causes an increase in the resulting flamespread rate. There are several mechanisms which operate to modify therate of flame spread. For example, (1) increased heat transfer tounignited surfaces by heated positive ions which move from flame zone tounignited surface of negative polarity; (2) increased heat transfer tounignited surface by the drag of heated, nomcharged particles along withheated positive ions as described in (l); (3) increase in localizedchemical reaction at the negative electrode by altering equilibriumconcentration of species; (4) control dissipation of electrical energyinto the surface of unignited propellant surface due to electronicconduction along propellant surface.

The power supply is connected externally and can be disconnectedimmediately upon ignition by means of quick disconnect mechanisms. Thiseliminates the need to mount a large power supply on board the missileand permits the power supply to be ground mounted or remain with aflying launch platform.

The power supply is a high voltage, low amperage unit which operates forthe very short time interval only, and is associated with ignition delayof the rocket motor.

Many variations of design and configuration are possible which willenhance flame spread. Reversal of polarity will cause similar but lessuseful effects. The changes in electrode geometry, which are possible,give greater flexibility to the system. However modified, the governingprinciple employed in this ignition system is the controlled movement ofheated ions and electrons from the flame zone along the unignitedpropellant surface, using electrostatic field forces developed withinthe rocket motor by means of an externally mounted power supply.

These and other advantages, features and objects of the invention willbecome more apparent from the following description taken in connectionwith the illustrative embodiment in the accompanying drawing.

DESCRIPTION OF THE DRAWING FIG. 1 is a schematic longitudinal sectionalview of a rocket with the device of the invention installed therein; and

FIG. 2 is a schematic cross-section taken on the line 2-2 of FIG. 1.

DESCRIPTION OF THE PREFERRED EMBODIMENT An internally ported solidpropellant rocket body is represented generally by the numeral 10 andthe nozzle by the numeral 1 1. r

The ignition system comprises a resistance wire initiator l2 installedat the forward end. It is to be understood that the invention is notlimited to the specific geometry shown. The showing of the initiator isexemplary only. Other more sophisticated initiators can be employed.

A resistance wire 12 is described herein only for purposes ofdiscussion.

The connections for the resistance 12 are made through forward adapter16 which may be made of high dielectric material such, for example, asmicarta which is phenolic or epoxy impregnated linen.

The resistance wire circuit is consumable and is vaporized immediatelyafter pyrotechnic material 1 has been ignited.

A high voltage is then applied between forward electrode 20 and aftelectrodes 22. The aft electrodes 22 are positioned on or near thesurface of the propellant grain l8.

It has been determined that the aft electrode wires have no observableeffect upon the flame spread rate in the absence of an electrostaticfield.

An electrostatic field is then produced between electrodes and 22 whichcauses a force to be exerted on certain charged particles (positiveions, negative ions and electrons) within the flame zone.

The resulting motion of these charged particles is primarily dependentupon the magnitude and direction of the applied field force and can becontrolled by external electrical circuitry. Movement of heated ions andelectrons from the flame zone along the surface of the unignitedpropellant causes an increase in the resulting flame spread rate by oneor more of several mechanisms. For example, (1) increased heat transferto unignited surface of negative polarity; (2) increased heat transferto unignited surface by drag of heated, non-charged particles along withheated positive ions as noted above; and (3) increase in localizedchemical reaction at the negative electrode by altering equilibriumconcentration of species; (4) control dissipation of electrical energyinto the surface of unignited propellant due to electronic conductionalong propellant surface.

The power supply 24 is connected externally and can be disconnectedimmediately upon ignition by means of quick disconnect 26. Thiseliminates need to mount a large power supply on board the missile andpermits supply to be ground mounted or remain with a flying launchplatform.

The power supply 24, indicated in FIG. I, is a high voltage, lowamperage unit which operates for the very short time interval associatedwith ignition delay of the rocket motor.

Although the invention has been described with reference to a particularembodiment, it will be understood to those skilled in the art that theinvention is capable of a variety of alternative embodiments within thespirit and scope of the appended claims.

lclaim:

1. In a solid propellant rocket motor ignition system, means for varyingand controlling the ignition delay time thereof comprising a hollowcylindrical propellant grain having an exposed internal surface, aplurality of aft electrodes positioned along the exposed surface of saidpropellant grain and extending forward from the rearward end thereof, aforward electrode positioned in the central forwardmost end of saidrocket motor and extending rearwardly along the central axis of saidrocket motor, a pyrotechnic material positioned around the base of saidforward electrode, initiator means in contact with said pyrotechnicmaterial for producing ignition thereof, and an electrical power supplyoperatively connected between said forward and aft electrodes forapplying an electrostatic field therebetween and causing a force to beexerted on charged particles within the flame zone thereby producingmotion of the charged particles proportional to the magnitude anddirection of the applied field force resulting in a correspondingvariation in the rate of flame spread along the surface of thepropellant grain.

